Oxidation and corrosion protective coating development for superalloys and refractory metals has been spurred by advances in propulsion technology in turbine parts since about 1960. These advances have placed increasing temperature and structural demands on materials for service at temperatures of 1010.degree. C. and above. Consequently, coatings are relied on to protect superalloy components such as turbine blades and vanes from environmental attack and to provide thermal barriers at the operating temperature of the superalloy component.
Improvements in the efficiency of gas turbine engines can in general best be achieved directly or indirectly by an increase in the temperature of the combustion gases incident on the turbine blades. The main constraint to the achievement of this objective is the limited choice of materials for the blades which will retain adequate strength and corrosion resistance above 1100.degree. C. for sufficient lengths of time. New processing developments for advanced nickel-base and cobalt-base superalloys have given the engine designer new limits of strength capability at the expense of environmental corrosion resistance. Simultaneous advances in coating technology have gone some way in achieving a satisfactory balance of materials requirements. However, further increases in gas temperature up to and even beyond 1600.degree. C. are still required. To meet this problem refractory alloys and ceramics must be considered as potential materials for advanced engines and progress towards reducing metal temperature is desired.
The principle of applying a low thermal conductivity ceramic, a thermal barrier coating, to a metal substrate as a means of thermal insulation has been recognized for some time. Many of the problems which have arisen in the past have been associated with metal substrate/ceramic compatibility. Differences in thermal expansion between the alloy and oxide invariably cause spallation of the thermal barrier layer. Adhesion of the ceramic composition to the substrate has posed further problems. Many of these initial limitations have been overcome by applying to the substrate a first so-called bond coat, e.g. of molybdenum, nickel-chrome, or MCrAlY, where M is nickel, cobalt, iron or mixtures thereof, followed by the preferred refractory oxide barrier layer, usually comprising some form of stabilized zirconia. Zirconia stabilized with either calcia, hafnia, magnesia, yttria, or any of the rare earth oxides may be used as a barrier oxide due to its very low thermal conductivity, low density and high melting point.
Engines for commercial aircraft, some military aircraft, and power generation service that have thermal barrier coatings eventually crack, spall, or undergo chemical and physical attack during their service life. Overhauls of these coatings are usually done periodically. During overhaul, turbine blades and vanes that have not exceeded creep limits and are not otherwise severely eroded or damaged are refurbished for reuse. Coatings, such as thermal barrier coatings and bond coats, are stripped from the components. The components are reworked and cleaned as necessary, recoated, and returned to service.
The thermal barrier coating repair on jet engine or power generation parts involves complete removal of thermal barrier coatings before recoating the surfaces with fresh thermal barrier coating and bond coat. The grit blasting method currently used to remove thermal barrier coatings is a labor intensive and time consuming process. In addition, it damages the bond coat as well, so that both the thermal barrier coating and the bond coat need to be refurbished. Also, repeated removal of bond coats thins the walls of the airfoils and increases the hole sizes in multihole blades thus increasing the airflow through the blades. As a result, only one full strip is allowed for repairing blades. Thus, there is a need to provide a process to remove thermal barrier coatings from parts without attacking or damaging the underlying bond coat or substrate.